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Why were most of the first-generation-Learjet wing mods never approved for the Learjet 23?
First-generation (straight-wing1) Learjets (the 23, 24, 25, 35, and 36) are, rather infamously, beasts to fly, due in part to their hideous stall behaviour;2 due to their flow-separation-prone laminar-flow wings, they have a tendency to stall at a fairly high speed with little to no warning and roll uncontrollably to one side, something that caused a considerable number of fatal crashes upon landing.
For this reason (and others), a number of modifications were devised over the years for the first-generation Learjet wing to improve its flight characteristics:

The Dee Howard-Raisbeck Mk. II wing mod, which lowered the aircraft's stall speed considerably, was released in 1975.
The following year, Gates Learjet responded with the Century III modification pack, which both lowered the aircraft's stall speed (by modifying the cross-section of the wing's leading edge) and modified the aircraft's stall-warning system to alarm earlier if the aircraft's angle of attack is increasing rapidly.
In 1979, the Softflite modpack (a set of stall strips and wing fences installed on the upper surfaces of the wings) became available; it considerably improved the straight-wing Learjets' stall characteristics, changing their behaviour from an abrupt stall and violent rolloff to a gradual, wings-level stall.

Learjet 23 owners, however, are unable to benefit from most of these improvements, as most of them were never approved for use on the 23.  The Mk. II modification was approved for the Learjet 23, but the Century III and Softflite modpacks - with the former's improved stall-warning system and the latter's benignified stall characteristics - were made available only for the Learjet 24 and up:

The manufacturer's Century III and Softflite wing modifications to improve the airplane's slow speed and stall characteristics have not been approved for the [Learjet] Model 23.  The Dee Howard-Raisbeck, Mark II, a similar wing improvement modification, has been approved for the Model 23, but the accident airplane was not equipped with this modification.  ...  [NTSB AAR-83-01, page 16 (numbered as 13).]

Why was the Learjet 23 excluded from most of these stall-protective modifications?

1: Technically, the wings of first-generation Learjets do have some sweep.  However, the wings' sweep angle is too small to make a meaningful improvement in the aircraft's flight characteristics; the 23/4/5 have the leading edge swept at 13° and the trailing edge unswept (giving a quarter-chord sweep angle of only 9.75°), and the wings of the 35/6 are based on the 23/4/5 wing, with the addition of completely-unswept wingtip extension panels which give it an even lower effective sweep angle.
2: Other bad points: an original autopilot with a great tendency to produce large, spontaneous pitch inputs in either direction, a tendency to dutch-roll violently if said autopilot failed or was disconnected at cruise speed, or if the aircraft's yaw damper failed or was disconnected at any time except when configured for takeoff or landing, a mach-tuck susceptibility best described as "extreme" at not far above MMO on an aircraft with easily enough engine power to exceed that limit in level flight (aggravated by an overmach-recovery procedure which, as it turned out, was very likely to send the Learjet into an unrecoverable dive), et cetera, et cetera.

The cost of approving an STC for an aircraft can be quite expensive and the return on investment may not have been worth it.
------------------------------------------
|Aircraft   |  # Built |  Year Certified |
------------------------------------------
|Learjet 23 |    104   |       1964      |
|Learjet 24 |    259   |   1966 to 1976  | (All variants)
|Learjet 25 |    369   |   1967 to 1976  | (all variants)
|Learjet 35 |    670   |   1974 to 1976  | (all variants)
|Learjet 36 |     64   |   1974 to 1976  | (all variants)
------------------------------------------

The Lear 36 is, for all practical purposes, a variant of the Lear 35.  If you take a look at the list, the Lear 23 is the oldest with the least number built and that would weigh heavily on whether or not to pursue an STC.  If an owner of a Lear 23 wanted to purchase the STC and absorb the cost of the certification, these would have been made available to those aircraft as well.
An owner can effectively upgrade their Lear 23 to a Lear 24 by completing and engineering change record (ECR 227, ECR 230 or ECR 233 depending on serial number).  Once the new Lear 23/24 name plate is installed, the STCs may be applicable.
It is also just a guess.  The FAA (as we know it now) was formed in 1967.  It is possible, the STC holders tried to get their modifications approved by the FAA for the Lear 23 but since the aircraft was certified before the FAA was formed, they may have declined or required the aircraft to meet current aircraft design standards and that may have been cost prohibitive.


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Why does the DC-9-80/MD-80 have this cusp in its fuselage?
The fuselage of the McDonnell Douglas DC-9-80 (also known as the "MD-80") has a pronounced cusp at floor level, as can be seen in this diagram (courtesy of Boeing, eater of McDonnell Douglas - at least, according to @JWalters):

Although the DC-9-80/MD-80's fuselage cusp is nowhere near as extreme as the Boeing 377 Stratocruiser's, it still seems like an unnecessary stress concentrator with no apparent reason for being there; why did McDonnell Douglas put in the cusp, despite its potential for accelerating fatigue damage to the fuselage?

Not a unique feature of the DC-9. It's easier to name mainline jet the aircraft that aren't double lobe so they don't have a cusp: A300/310/330/340 (same circular fuselage), A380 (elliptical), 777 (circular).

It was inherited from the DC-8. The DC-8 was initially planned with a smaller fuselage width but the upper half was enlarged when the Boeing 707 was made wider before it. The DC-9 cross section was a smaller version of the DC-8 cross section.

It's not a stress concentrator; it's just the opposite.  What you're missing is that the floor itself at the pinched part forms a tension bridge that allows a more or less 'ovalized' circle while still maintaining tension loading on the skins and frames as if it was a pure circle.
If I had a rubber balloon filled with air and was able to run a string internally from one side to the other, attached to the walls of the balloon, and then drew the string in to pinch the sides of the balloon into a figure eightish profile, I'd have the same thing.  All the loads on the skins are still in tension, as well as the bridging floor beams (the string).

Bubbles seem to demonstrate that this is the most stable shape: https://goo.gl/images/82C3br

To add to John's answer, why did they do it?  To make more room in the passenger area.
It is important to remember that, opposite a submarine, pressure is higher on the inside of the aircraft in flight, rather than lower.  So structural design favors a cross brace to hold it together.  A submarine would be strongest if it were perfectly spherical.
Also, notice the "pinched" fuselage has lower frontal area, resulting in better fuel economy from less drag.
The MD-80 design, with its clean wings (engines mounted on rear fuselage), lived on as the
Boeing 717.

The passenger area would be larger if they removed the concave area, so I don't understand the point of your first paragraph.

Could be statisticly passenger + seat has more volume on average than luggage. But, apparently, enlarging the upper half was the best engineering choice.  Also seen on the Super Guppy.

I mean using an egg-shaped cross-section. You see the point where it pinches in? Push that out to give a smooth curve: the passenger area gets larger and the floor still acts as a tension beam.

@David Richerby: Shoulders are generally wider than feet, and that's where the increased room is.

@David Richerby good thought.  It looks like they needed to fit the 5th row of seats in onto an old "chassis" to save $.  The end seats look a little tight.  I believe the 717 is no longer in production.

@DavidRicherby An oval shape like that disrupts the circular arc for each part circle and introduces bending forces in the frames above and below the floor beams as pressurization forces push out.  You can do it that way, but you will add a bit of weight stiffening the frames.  The double bubble profile is the lightest way to achieve an oval-ish cross section.

@RobertDiGiovanni: No, the DC-9 had five-abreast seating going all the way back to the DC-9-10 (entered service 1965).

I would say the opposite is true: it increases room in the cargo hold, not the passenger area.

This is called the "double bubble". The 737 has a similar but less pronounced design.
Both aircraft had a primary requirement to seat a specific number of passengers in each row: 5 in the DC-9 and its successors, 6 in the 737. Seat-bottoms are the widest parts of the seat, so they made the fuselage the widest there.
But continuing arc downward from there would result in insufficiently shallow cargo area, so they designed a lower bubble that provides more width and depth in the hold. And they net out to less materials (weight) and cross-section area (drag) than if you tried to fit that many seats per row in a purely circular cross-section.

The answer is soap bubbles:

They are filled with slightly higher pressure air, and when they're attached to each other, they have a planar "reinforcement-like" part between them. (see the picture)
This is similar to your "attached circles with reinforcement between them"-style structure. (DC fuselage)

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Why doesn't extending the flaps on a B-2 send it into an uncontrollable dive?
The Northrop-Grumman B-2, a tailless-flying-delta-wing aircraft, has a set of flaps along its trailing edge, as seen (for instance) in this cutaway diagram:

(Image from Funker530.)
But tailless delta-winged aircraft (except for those few equipped with canards) can't generally use flaps, because a tailless canardless delta-winged aircraft has no way of countering the nosedown pitching moment produced by extended flaps.
How does the B-2 avoid pitching down uncontrollably when its flaps are extended?

Doesn’t it simply cheat on the „tail-less“ part by deflecting the tail?

Other sources identify these surfaces as elevons. Are you sure they are actually flaps?

@Sanchises They are listed as elevons in How Stuff Works, with the "beaver tail" as a Gust Load Alleviation System.  The B2 Spirit has clamshells at the trailing edges of its wing tips, which open top and bottom as airbrakes.  These may act as "flaps" when used together, or "rudders" when used on side at a time.

The Aeriane Swift ( https://en.wikipedia.org/wiki/A%C3%A9riane_Swift ) is a tailless aircraft that uses flaps.  But, they appear to be located near the CG.

I don't think the flaps are used as high lift devices. Watch this video of a B-2 landing. Perfect view:

I am sure the B-2 is an unstable aircraft meaning if it pitches up the pitching moment increases and the aircraft wants to further increase pitch.  This is contrary to traditional design, but can be controlled through fly-by-wire. This increases the efficiency of a flying wing design as many of the things done to help flying wing stability hurt efficiency.
This means that when the flaps at the trailing edge deflect they are helping stabilize the aircraft.  This means though that the angle of attack or relative angle of the wing to the incoming airflow cannot go negative.  This is not a problem as the aircraft will always be pitched up with flaps deployed.
I am an aerospace engineer who is currently working on flying wing type aircraft.

"This means that when the flaps at the trailing edge deflect they are helping stabilize the aircraft.: -- does this really follow from your earlier comments?

First we consider the purpose of a "tail", more correctly Horizontal stabilizer, for pitch control.  The classical set up for a straight wing is to have a tail for adequate pitch stability.
Proper design, especially for a bomber, calls for pitch stability to be stronger than all but the strongest control inputs, including flaps.  You don't want to stall one of these.
The B-2 has enough pitch stability in its design not to need an extra tail sticking out the back.
Any pitch down moment caused by flap deployment is not only counteracted by the aerodynamic stability of the aircraft, but also computer control of surfaces such as elevons and beaver tail.
The "tail" is hidden in the structure and the computer control.  Finding the "tail" leads to
an understanding of how it works.  Swept wings lengthen the pitch axis just like a fuselage.
Surface area at the trailing edge of a delta serves the same purpose as a horizontal stabilizer.  Deltas are really blended wing-tails.

Since the trailing edge of the B-2 wing-- ignoring the large serrations-- is essentially a straight line running in the spanwise direction, it's unclear how deflecting the elevons would be an effective way to counter the nose-down pitching moment generated by deflecting any flaps that may be present.  The elevons are operating at no greater a moment-arm, in the fore-and-aft sense, than the flaps are,

"Proper design, especially for a bomber, calls for pitch stability to be stronger than all but the strongest control inputs, including flaps. You don't want to stall one of these."-- is this really an appropriate statement in the context of an aircraft with artificial computer-generated stability?

Hey, long time no see.  Please look at the above picture.  They use clamshells to achieve the same effect as flap down elevon (or spoiler) up.  And with dual rates, one can increase elevon (or elevator) throw if needed.

Model B2s fly without computer control. The statement is appropriate as a design goal, whether it is achieved by passive aerodynamic stabilization (old fashioned) or by computer control, or by BOTH. And those serrations, in addition to being stealthy, provide MORE than a straight line of space for control surfaces. But no, they don't have traditional flaps, the diagram in the question was mis-labeled.

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How can seagulls slide in the wind without flapping their wings?
There are a lot of seagulls in a park near where I live. Sometimes I can see them somehow "sliding" forward or backward diagonally when flying against the wind, however, their wings aren't flapping. What are the flight mechanics behind this?
Can gliders mimic this style of flight?

I've never heard of other birds than albatrosses do dynamic soaring, but since this is over land, it can be simpler ridge soaring—as the sea breezes hits the shore (and obstacles on it), it turns upwards and the birds simply fly in this updraft. They are flying against the wind and gliding down relative to it, but when you add the speed of the wind, they may be stationary relative to the ground—or zig-zag from side to side if the wind speed is a bit less than the forward speed the birds need to efficiently glide.
Gliders can do the same, but since they are bigger, they need a mountain range where the seagulls can do with a several feet embankment.

I couldn't find the dissertation itself, but there is apparently a dissertation on seagulls dynamic soaring, by Dr. Ferdinand Hendriks 1972 at UCLA. Edit: this paper references Hendriks 1972 https://cfwebprod.sandia.gov/cfdocs/CompResearch/docs/02-1896_MobileSensorArrays.pdf

@AlecMartin, interesting. The description in question does not really match dynamic soaring anyway though, since dynamic soaring involves flying circles, with low and high part when using the boundary layer of the ground.

If your perception is that the seagulls are pointing straight into the wind but moving sideways to the wind, I suspect that you are mistaken.  When the wind is strong and the groundspeed is very low, it only takes a slight difference between aircraft (or bird) heading and wind direction to produce a large crosswind component in the ground track.  This is certainly true for gliders (and airplanes) as well as for birds.
So "sliding sideways" in relation to the wind is accomplished by steering the aircraft (or bird) to point slightly to the left or right of where the wind is coming from.  In strong wind, an aircraft or bird can move perpendicular to the wind direction, with no "forward" progress, while still pointing almost directly into the wind with wings level.  Note that this does not actually involve any sideways component to the airflow over the aircraft (or bird).
If the seagulls you observed are maintaining altitude while doing this, without flapping their wings, then they are undoubtedly in slope lift.  Slope-soaring on a windy day often involves pointing almost straight into the wind, even while the ground track follows the line of the slope, which may be perpendicular to the wind.

Some birds 'harvest' energy from the wind gradient. Perhaps it's the case of the seagulls that you have observed.
https://en.wikipedia.org/wiki/Dynamic_soaring

I'm sure the more probable answer is ridge lift.

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How does the nose wheel tire stop rotating on retraction as the aircraft takes off?
As we know, the nose wheel tire and main gear tires are rotating when aircraft is taking off and the wheels are retracting. Thus there should be a mechanism to stop this rotation in order to prevent oscillations transmitting to fuselage and damages to wheel well so auto brake system is used in main landing gear brake system to counter this problem. What mechanism used to stop rotation the nose wheel tire? Is it a shoe brake system?

Since the question references the nose wheel specifically, this is not an exact duplicate of the linked question. The answer is that there are brakes in the form a pads attached to ceiling or front wall of the wheel well. From the 757 AMM "Two wheel spin brakes, attached to the nose wheel well ceiling, stop the nose wheels from spinning after takeoff and gear retraction. The brake pads are installed on the contact surfaces of the spin brakes to provide good frictional contact."

@SportsRacer: Good comment, voting to reopen, so that you can provide a complete answer.

Voting to leave closed. The other question covers both main and nose gear, and DeltaLima's answer to it already mentions the pads in the nose wheel well. If Sports Racer wishes to post an answer, s/he can do it there.

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How do fighter jets measure how high they are?

I know most of the commercial planes use altimeter to find how high they are. I was just curious in finding out what fighter jets use to to find their height and what are the other technologies that could be used other than pitot tubes to find speed of an aircraft or fighter jet.

2. What material are used to make pitot tubes in aeroplane?

Why do you think military aircraft are significantly different in this regard to commercial planes?

Also, please ask only one question in each question post. The expertise needed to answer your two sub-questions is not the same, and they touch on very different subject areas, so it's best to ask them as two separate questions. (The fact that you felt the need to separate them is a hint that they are too different to fit within a single question.) Questions that ask multiple separate questions at once are liable to being closed as "too broad". You can avoid that by making sure to only ask a single question per question post.

Fighter jets have the same system as commercial aircraft: An altimeter based upon external air pressure. this measures their altitude relative to pre-defined atmospheric conditions. Many will also include a secondary system that relies on GPS (or local equivalent) signals to produce a height above the reference geoid rather than an atmospheric datum.
Regarding your second question - you should only post one question per topic. An answer is difficult as the construction varies significantly between manufacturers. Many pitots are made of steel variants, however.


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Can the four-engined Lockheed JetStar II fly on one engine?
As we know, there are no four-engined business/private jets currently in production. Some Dassault Falcon models are actually trijets, but this is not the case. The only four-engined business jet in serial production was Lockheed JetStar line.
I'm wondering, in case of extremely rare emergency event, when 3 out of 4 16.5kN TFE7313 engines shutdown, could JetStar constantly fly on some flight level without loss of altitude using only one engine?

I'm pretty sure the answer to this is no. Even most twins can't hold altitude on a single engine. There is an old saying that when you have a multi-engine aircraft with an engine out, the other ones just get you to the crash site faster (or something like that).

Possible. But at a very high power setting for that engine and a healthy boot full of rudder to track straight. And your best rate of climb will be anemic, if any at all.

@RonBeyer  ALL twin jets not only CAN hold altitude single engine, they can climb out on 1 at max t/o weight -- it's a certification requirement. The real answer to this question can only come from the JetStart performance book, but it wouldn't shock me if the answer is, on a cold day with a light load, yes. I could also believe "no" -- but it will take the performance charts to resolve this. A sufficiently light 777 almost certainly CAN hold level flight on 1/2 the rated thrust of 1 engine (1/4 full thrust), so it's not unimaginable that a JetStar might also, under the right conditions.

@RalphJ Complete mind slip on my part, I was thinking of twin pistons. I'm still under the impression of it not being possible to hold altitude on a single engine of four. As you say, twin jets are designed to lose an engine, quad jets are not designed to lose three...

I have 3000 hours in jetstar.... The answer is no... It will not fly on one... Two yes and you really don't need that much rudder.. In the simulator with two out I used to put the inboard operating engine at 98% and modulate the outboard engine to control speed/sink doing that it really only needed 1/4 rudder trim

Thanks for chiming in with actual experience! Hope you stick around to provide more great answers.


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What brakes are used between line-up and takeoff?
I was wondering if an airliner was instructed to line-up on the runway ("line-up and wait"), will the park brake be used or will manual braking be used (pedals) during the wait until the take-off clearance is given? Same for holding short the runway before line-up / take-off clearance? Does it make a difference which brakes are used in the first place?

If you are on the runway it should be the toe brakes, it is typically a very short hold.

Somewhat Related: I some times get a lineup and wait in my little Piper. In that case toe brakes are used. On small planes the E-Brake usually just actuates the regular brake system anyway. As Ron mentions, you never really wait that long.

This question is like asking "Do you step on the brake pedal or set the parking brake while waiting for a red light when driving". It boils down to (1) how long do you expect to wait & (2) whether want to rest your feet at the expense of using your hands.

Just toe brakes. Typically the instruction between line-up-and-wait and clearing for takeoff is very short, maybe ten seconds.
In simpler aircraft, the parking brake just "pinches" or closes off the brake fluid line to keep the wheel brakes held, but in the fly-by-wire airliners, the parking brake just clamps the normal wheel brakes as if the toe brakes were held.

In addition, most companies (that I know of) forbid the use of parking brake on runway. That is to reduce the risk of trying to take-off with parking brake on. There have been some accidents because of that.


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What are the weights of vertical and horizontal stabilisers of a Boeing 747-400?
For my math exploration (IB) I am doing something related to aircraft. So I was wondering what the weights of the horizontal and vertical stabilisers were. If you do know the weights please state the source of the information too.

The wings are where fuel is stored, so their weight change depending on how much fuel is on board. Are you referring simply to the weight of the wing structure? By "the back ones", are you referring to the horizontal stabilizer?

I don't know if it helps, but the basic operating weight of the 13 747-400 freighters that I have current data for runs from a low of 351,088 lb to a high of 365,970 lb.

If you're trying to approximate the center of mass, the computation of the longitudinal component is required to be done before every flight of large aircraft as a proper CG (which is what the center of mass is) is critical to the safe operation of the flight. To get the longitudinal CG is simply a matter of taking the known CG of the empty aircraft and adding to it what you put in the aircraft. So, would that calculation satisfy the requirements of your math exploration?

For Boeing 747 aircraft, the information you're seeking can be found in the WEIGHT AND BALANCE CONTROL AND LOADING MANUAL that Boeing issues for each aircraft manufactured. See the COMPONENT WEIGHTS AND BALANCE ARMS starting in section 1-86-001. A number of these manuals, all out of date as I don't keep them up, are at http://terryliittschwager.com/wbmanuals.php.

From my time as an intern at the plant at Paine Field, I know that a 747-400 wing and center fuselage section weighs 42 tons.  That's just the structure itself.

From the table in this answer: the tail group of the 747-100 and 747-200 is 11,850 lbs. In order to find the weight of horizontal and vertical stabiliser, we can divide up the weight according to tailplane surface.
Table 9-2 in Torenbeek gives the ratio of horizontal tail surface/wing surface as 0.267 in table 9-2, for both the 100 and the 200. Wikipedia gives the wing area as 511 m$^2$, so horizontal tail area is 136 m$^2$.
In the same book table 9-3, vertical tail surface ratio is given as 0.196 so vertical tail area is 100 m$^2$.
Weight for the 747-100/200 is therefore:

Horizontal stabiliser = 136/236 * 11,850 = 6,830 lbs
Vertical stabiliser = 100/236 * 11,850 = 5,020 lbs

Above weights is with all equipment installed: actuators, hydraulic system lines, rudder & elevators, stabiliser incidence electric motors etc.


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What exactly does it mean, that plane has diverted due to the needs of a passenger?
Twitter account for Flightradar24 has reported twice in past 24 hours about two different British Airways flights that has diverted their original route "due to the needs of a passenger" (here and here).
What exactly does it mean? Is it:

a passenger requesting to land on other airport, because he forgot his wallet?
operational decision to put more passengers on-board empty plane?
something else?

More than likely its a medical need urgent enough to divert. No pilot is going to turn a plane around to get something a passenger left behind.

@RonBeyer unless it's a business jet and the passenger is the one chartering it :)

As suggested by @RonBeyer,  "needs of a passenger" is not "at the request of a passenger"

It seems, that I was confused by FR24. I'm watching their Twitter posts for about past 2-3 years and never observed "passenger needs" before. It was always "medical urgency", "operational needs" or "reason unknown". I wonder, does this fact really mean, that I've asked the wrong question? What is so wrong in trying to learn about something unknown? Strange...

@trejder you jsut discovered the I don`t know therefore downvote/opinionbased etc etc condescending bandwagon :-p What you could do is to email or twitter-ask the tracking service and find out! :)

You are asking for an exact meaning for a term meant to cover a variety of situations.

@KorvinStarmast Your comment stands in an exact opposition to Ron Beyer's comments which clearly states that "needs of passenger" term covers exactly one situation -- medical reasons. Your comment also stands in opposition to most other comments and answers given here that claims the same.

Your question asking for "exactly" is the problem I am identifying for you.  Words have meanings.  you can either improve your question or not.  I don't make or lose any money either way.

The only passenger needs that would result in a non-scheduled landing would be a medical emergency or acts/threats of violence of the level that prompt removal of the passenger is needed (with law enforcement coordination).
A passenger forgetting his wallet would not result in a diversion.  An aircraft deviation to enplane more passengers and proceeding to the scheduled destination would be a highly unusual situation and would not be for passenger needs, it would be for operational needs.

Thank you for your answer. The "missing wallet" was an obvious joke. It seems, that (too) many people misunderstood my motives and taken in (too) seriously.

I don't think they would call threats of violence "Needs of the passenger." Almost certainly health.

@user3344003: That could be considered as the passenger having a need to be arrested.

The transportation industry tends to use specific phrases to describe events of this kind. They are not just euphemisms, but are intended to neutralise the potential for drama, speculation, voyeurism that can surround them.
In the UK, for example, you will sometimes hear that an London Underground service has been disrupted by "passenger action". This means that someone has committed suicide on the rail lines.
As noted in other answers, "the needs of a passenger" means a serious medical issue.

Dutch Railways use "collision with a person".

"someone has committed suicide on the rail lines" -- not necessarily; I think they use it for any time a person ends up on the track which (a) might not be fatal; (b) might not be deliberate (drunken stumbling; pushed). I'm not entirely sure the person has to end up under a train.

@RogerLipscombe You are probably right.

According to FR24's comment on the post, they use that term when they do not know the exact nature of the incident. However, it is most likely a medical emergency.


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How does the safety record of the twin piston Diamond DA-42 compare to the single piston DA-40?
Its fairly well known that the Diamond DA-40 has one of the smallest fatality rates of any GA plane. The Diamond DA-42 is a twin engine aircraft with a very similar design to the DA-40.
Now, it has been shown that, in general, twin aircraft are not that much safer than single because the added complexity of operation removes the added safety from engine failure. However, the DA-42 is much easier to fly with FADEC control of the engines and a lot of automation.
Is the DA-42 safer than the DA-40?

A huge amount has been written on the topic of single vs. twin. Here may be a start: http://www.flyingmag.com/wrong-worry-twins-versus-singles

Be careful making any relational inferences.  I watched a news broadcast once where the reporter, straight faced and without the slightest hint of irony or understanding, proclaimed that twin engine aircraft were twice as likely as single engine aircraft to experience an engine failure.  And he just left it there, with no further explanation, for the ignorant masses to digest…

So it appears that Diamond answered this question for me.
See this page:
http://www.diamondaircraft.com/about-diamond/safety-features/

From this chart, it would appear that the DA42 has a slightly higher accident rate, and of those accidents, even more are fatal (which makes sense due to the higher speeds). Still loads safer that the SR22 though, and I'd personally prefer the 42 over water as well.

The difference in accident rates between the DA40 and DA42 is not necessarily a function of how many engines the aircraft is equipped with.
There are numerous possible causal factors for aircraft accidents which could be responsible for the difference in accident rates between these two types.
The first thing that comes to mind is that the DA40 is very difficult to stall.
When demonstrating stalls as a flight instructor you have to crisply and deliberately pull the nose up to get a stall "break". If you are not aggressive with the elevator, and you pitch up gradually to maintain altitude as speed bleeds off (like you do in aluminum airplanes) it wont give you a stall break- it just mushes and loses altitude at its low stall speed of 49 KIAS (flaps down).
While demonstrating a "Power On"/"Departure Stall" with full power, it is nearly impossible to get a stall break unless you are extremely aggressive.
Thanks to its low wing loading, the DA40 can be landed somewhat slowly and, perhaps, even softly in any random field or open space and have a relatively short landing roll.
This is not the case with the heavier DA42.
In addition to being a light aircraft with a high lift wing and excellent, low speed, climb performance, the DA40 is unbelievably easy to fly.
The DA42 is a heavy, faster aircraft with higher wing loading and a (marginally) higher stall speed of 57 KIAS (flaps down).
Unlike the DA40 you don't have to force it as enthusiastically to get it to stall. It is a heavy airplane. It will stall.
Unlike the DA40, the DA42 takes FOREVER to accelerate to takeoff speed and requires a lot more runway (1400 feet vs the DA40's 800 feet, depending on temperature and Density Altitude). This is not only because it is heavy, but because the TAE 125 engines are only 135 horsepower each. The only reason the DA42 can cruise at 155 KIAS is because it is so aerodynamically clean (unlike a Piper Seminole which requires 180 horsepower engines to cruise at the same speed).
If most accidents happen during takeoff or landing, this unusually low power output is not an advantage for the DA42 (nor is the retractable gear which could be forgotten).
In comparison, not only is the DA40 super easy to fly and hard to stall, but the engine responds immediately with a good amount of power (in comparison to its weight) and it climbs like a rocket and can easily clear obstacles after a short takeoff roll.
Additionally, the DA42 takes A TON of right rudder to keep the nose on centerline on takeoff (enough to make your calf sore).
If you lose directional control and go off the runway you just became one of those accidents in the statistics (another one that would not happen to a DA40).
DA42s are used for multi-engine training. Unlike with the DA40, there is a possibility (though slight) of an accident related to engine-out training.
Although a simulated engine-out landing in a DA42 is a "non-event" if you have the rudder trimmed correctly (and you do have the simulated dead engine available to use if you had a real need for it).
With fixed landing gear, the DA40 cannot be involved in a gear-up landing, which could contribute to the DA42's higher accident numbers (especially since so many are used for multi-engine training).
Perhaps an even more likely cause of the DA42's accident rate is the type of mission it is used for...  while a DA40 will often be used for fun relatively short VFR flights where no weather is encountered, it is more likely that a DA42 is going to be used for cross-country "mission"-type operations where you need to get somewhere, making it possible for the pilot to experience "get-there-itis" which causes the pilot to continue into weather they should not.
These longer "mission" type flights will also take the aircraft into environments where you may encounter weather that is not expected or is possibly hazardous (depending upon the accuracy of the weather forecast at the time of departure).
If the DA42 is equipped with the TKS anti-icing system the aircraft is far more likely to be used on a flight where icing may be encountered, which introduces weather conditions/hazards that the DA40 is far less likely to experience, due to not being taken intentionally into icing.
Depending upon the attitude of the pilot, the TKS anti-ice system could give a false sense of confidence to enter icing conditions which should be avoided.
Note that for the TKS to work you must coat the wing with the fluid BEFORE you enter icing conditions, not after.
Or perhaps the difference is all due to the takeoff roll taking far far longer than anyone expects...  And gear-up landings.
We had a customer many years ago who landed one of our DA42s with the gear up.
Interestingly the aircraft sat on it's exhaust pipes at the front and on the ground-down boarding steps at the back, the structure did not touch the ground anywhere. Of course the props were destroyed.
Regarding twins being statistically "less safe":  A twin can be more safe if the pilot is is mentally "spring-loaded" for an engine failure on takeoff, proficient at engine failure procedures and single-engine flying, and calculates the single-engine climb rate for that day's conditions to ensure that all obstacles can be cleared.
Of course this is dependent upon promptly and properly responding to an engine failure calmly to identify the failed engine correctly, establish zero side-slip and clean up drag (flaps, gear, feather prop).
Sadly, the statistics show that many multi-engine pilots are not ready for an engine failure and react with surprise and lack of proficiency. With the DA42's FADEC (no mixture or prop controls), it could not be easier to get the airplane into the correct configuration.

I know its an old post, but I would add that to me its important to factor in the speed of the aircraft.  Speed as in how far are you going to travel in an hour.  If you factor in that for the same distance flown between points, the DA42 will get you there faster, then both aircraft come in at about the same accident and fatality rate.
It's not a perfect analysis, but it makes sense.  It's not that the DA42 is more dangerous because the twin is faster, its simply that to accomplish the same mission you are going to be in the air for fewer hours in the DA42.

I would note how this is not in favor of the DA42. If you have more accidents in less flight hours, your aircraft is less safe.

@Federico, it depends on your goal: if the point is to get from A to B (as it is in most cases, esp. for DA42), then the proper metric is accidents per distance (km or mile). Here DA42 indeed may have an edge, even with higher per hour rate. (Needs calculation). For sightseeing, accidents/hour may be more representative.


text:
Can a pilot take off at his own discretion?
[I'm referring to an ATC conversation during the military coup in Turkey with an MD83 EP-TAS (ATA TBZ5402).
Can a pilot take off at his own discretion?
To summarize the link: a pilot is begging the tower to be allowed to depart and repeatedly denied due to what the controller describes as a "military operation" or "people in the tower". The pilot indicates the situation is dangerous and even reports seeing gunfire on the ground at one point. The pilot then repeatedly asks for permission to depart "at my own discretion". In the end, it appears that he does not depart.
Would there be any consequences for the pilot? Is he allowed to take off if the situation on the ground is more threatening than in the air? Are there any international regulations in this case?

In the US, there is 14 CFR 91.123(b) which states that a pilot may deviate from any ATC instruction in an emergency. I don't think there is an international regulation that covers a pilot in all countries though, and I'd hazard to say in some countries deviation may be met with military action.

Yes, it is a world wide rule that pilots can deviate from ATC instructions if failing to do so will endanger the aircraft. (Why on earth would the US be the only ones to have such rule!?). The question of course is how authorities interpret whether the actions taken did indeed prevent a dangerous situation.

Summarize the link, rather than just linking to it.

To summarize the link... a pilot is begging the tower to be allowed to depart and repeatedly denied due to what the controller describes as a "military operation" or "people in the tower"... The pilot indicates the situation is dangerous and even reports seeing gunfire on the ground at one point... the pilot repeatedly asks for permission to depart "at my own discretion"... in the end it appears that he does not depart (much to the disappointment of my desire for drama!)

Related question... does anyone know if there were in fact aircraft that departed IST without clearance?

If a pilot believes that staying on the ground is likely to result in the consequence of death, what would he have to lose by taking off?

@WBT: well, if there are two such pilots at the same airport at the same time, they might be exchanging the likelihood of death for the certainty of death. Presumably this pilot never became confident that the danger on the ground was definitely greater than the danger of going against ATC instructions. Which in turn is greater than the danger once he gets into the air. Hence his strong desire to get clearance.

@SteveJessop That airport has two parallel runways so not necessarily.  Would pilots be likely able to see each other out there and not cross/enter a runway when they can see another airplane in position for takeoff, similar to how surface vehicle drivers avoid collisions?  If not, can their radios help coordinate?  Obviously, an authorized/cleared takeoff is better, but I think "certainty of death" is overstating things even if there were two pilots reaching the same conclusion at the same time.

"Would pilots be likely able to see each other" -- they have a fair chance, but of course there have been collisions between aircraft on runways before, and that's even with ATC oversight. My point is just that it's the pilot's responsibility to weigh the risks, and there are things that he has to potentially lose from taking off, even given there is danger on the ground.

... for another risk, I don't know whether a pilot on the ground in a big jet can tell if there's anyone on approach to land. He doesn't have a rear view, but he might hear their comms. But in good visibility the approaching aircraft is presumably going to consider another aircraft on the runway to be a pretty sound argument for going around ;-) Obviously it's not "safe" in general for pilots to just do their own thing or we wouldn't bother with ATC at all. This pilot on this day had to decide whether it was "safer" than being on the ground, and decided it wasn't.

@SteveJessop You may want to consider that operations at non-towered airports happen thousands of times a day, even commercially without ATC assistance. Pilots are well trained to coordinate with each other and listen to the frequency for conflicting instructions. Yes conflicts do occur, but the majority of operations without ATC go ahead just fine.

@WBT In an armed insurrection, the risk of death on the ground has to be weighed against the risk of being shot down if you take off against orders to the contrary.

Generally speaking a pilot can (physically) do whatever they want: ATC doesn't have a remote control. If you push the throttles forward and pull back on the yoke you will take off, no matter what mean things the folks in the tower are shouting at you over the radio.
Practically speaking doing so can get you into trouble: You may be met in the air by planes with more guns than yours, or projectiles from the ground may be fired in such a way as to intersect your flight path ("You might get intercepted or shot down by very annoyed military types.") -- even if that doesn't happen you may face administrative or civil action for ignoring ATC instructions in the jurisdiction where you did this.

I'm not sure what the ATC policies are in Turkey, but the US FAA has a phrase specifically for this: UNABLE TO ISSUE DEPARTURE CLEARANCE. DEPARTURE WILL BE AT YOUR OWN RISK. (you'll find it in JO 7110.65W, on page 3-3-1).
The emergency authority of a pilot is also recognized worldwide (at least in ICAO states - you'll find it in Annex 2, 2.3.1, conveniently quoted on Wikipedia), so if the pilot in command deems that remaining on the ground constitutes an emergency they can violate regulations to the extent necessary to meet the emergency (i.e. forego getting a clearance and take off even if ATC is saying not to).

Helis often take off from taxiways and from non-movement areas, and can't be cleared. These are also "AT PILOTS RISK"

@rbp Indeed, if the area is not designated as a helipad takeoffs or landings from the ramp would be permitted "At your own risk" - the phraseology for that can be found in the same FAA order.

Except in extreme military situations such as in the example, taking off without clearance would never cause the dramatic response described here.  In the US it typically escalates from a warning and reprimand, to a helicopter intercept if necessary, then in rare circumstances they bring out the jets.  Even Adam Leon, a suicidal foreign pilot who consciously ignored intercept orders for hours, didn't have missiles fired at him.  See http://www.airspacemag.com/flight-today/dont-cross-that-line-5841988/?no-ist

@CodyP Bear in mind that the entire world is not the United States - We have a very relaxed system, but in many countries ATC is a military function which also happens to serve civilian traffic. An escalation to military response (interception, and yes - even being shot down) is entirely within the realm of possibility under such systems. Certainly more likely if there's an ongoing coup, but not something I would rule out with broad strokes.

Obviously, a pilot is physically capable of taking off if, e.g., the runway is clear. And, although the question uses the word "can", it's clearly asking if the pilot may take off.

exactly, it's totally clear what the use of "can" means in this question.

Yes, a pilot-in-command may take off at his own discretion if absolutely necessary in the interest of safety.
Rules of the Air (ICAO Annex 2) apply:

2.3.1 Responsibility of pilot-in-command
The  pilot-in-command  of  an  aircraft  shall,  whether  manipulating  the  controls  or  not,  be  responsible  for  the  operation  of the aircraft in accordance with the rules of the air, except that the  pilot-in-command  may  depart  from  these  rules  in  circumstances  that  render  such  departure  absolutely  necessary  in  the interests  of  safety.
2.4 Authority  of  pilot-in-command of an aircraft
The  pilot-in-command  of  an  aircraft  shall  have  final  authority as to the disposition of the aircraft while in command.


text:
What is the CDSU for a governor with magnetic trim?

The governor is a spring-loaded, flyweight-operated, hydraulic control valve whose driving gear is in mesh with the differential output ring gear through the scavenge pump gear. The governor senses CSDU output speed and controls the oil pressure to the control cylinder, maintaining a constant output speed of 12,000 revolutions per minute.
This I have understood from this picture but I want to know the operation in detail.
What is the purpose of the CSDU?

Please explain what the acronym CSDU means. Please add some background information on where in an aircraft it is found. Please add reference to your image: where did you find it?

sorry I just rephrased the question, the thing is in my book it is not a detailed explanation of how it works.

The CSDU function is to maintain the alternator constant speed by adjusting oil pressure when the engine RPM change. The flyweights are moving a swashplate not shown here. The magnetic coil is used to include the electric load into the regulation loop, as a high current will slow the alternator regardless of the engine speed. You can find the rest starting page 1-34 of Avionics In Plain English. Refine your question if needed after your have read this section.

thanks a lot, and will keep that in mind next time i have a question

I find it quite hard sometimes to understand procedures here. There is a question on an aviation related subject. The OP-er is asked to provide more detail, which he does. A concise answer is given in a comment, proving that the question is narrow enough to be answered by someone on this forum. And then the question is closed as too broad.

@Koyovis: The question first was "explain this mechanism", so the closure for being too broad. Then the question was refined, but for me still too broad. After the OP has a better understanding (e.g. by reading the section linked), if they are still missing something then they can edit and reopen, a request which I think will be supported by anybody and full answers could be provided. I believe this was managed correctly, and I agree this is a good topic (there was no need to downvote).

@mins Just wondering about the timing, the question was closed after your informative comment, which would have made a great answer by the way.

@abdulla - I've made a small edit that I think pulls out your main question. Please feel to clarify further from there if that's not what you're after. If this is what OP is after, then this is now a pretty focused question that can be reopened

Please reopen this question. After the edit it has a nice answer :-)

the question is open for comments and answer, @FreeMan, when I asked the question It was because the knowledge I had was limited and I wanted to know understand more about the diagram

and thanks for the answer @mins,


text:
Where can I find this transonic prop airfoil?

I am interested in identifying this prop airfoil due to its uncommon high .92% efficiency (the best metal props are closer to 83-87%). I have done an exhaustive search for it. It is labeled "NACA RML 9 G 06a". I do not know where I originally got this plot and therefore have no way to know what "ref 3" is.
It does not appear to be a airfoil number and a google search doe not return any "NACA RML" however I have a library of about 2000 "NACA RM" which refers to NACA Research Memos; perhaps the additional "L" means "Letter(s)".
I can find NACA reports starting with "RM G-10" but no G-9. I have also searched for derivatives of "06a" in an airfoil number. I would appreciate any assistance in identifying this airfoil.

The label you seek is not an airfoil, it is a research memorandum produced at Langley. (RM Lxxxx) The subject airfoil is NACA 4-(5)(08)-03 and the maximum efficiency from the test as reported in the conclusion was 88% at m0.71, partway down the curve where compressibility effects become adverse.

Thanks... wonderful! I see now the chart label was wrong and that explains why I could not find it.

@jwzumwalt What, if anything, did you wind up doing with the airfoil? It has interesting performance.

I am installing a VW engine in a home-built aircraft of my own design and was intrigued by the high speed efficiency. A VW engine is rated at 5000rpm but 3800-4000 (other builders typically run them at ~3400rpm) is more realistic and this puts a 55-60" prop near the speed of sound. This airfoil is about 5% better than similar designs at that speed.

@jwzumwalt How will you get the airfoil coordinates? It is not numbered like any NACA series with which I am familiar. Hope it works out for you.

I have pictures of it, and it appears to be similar in design and characteristics to a NACA 10-(5)(08)-03, NACA 4508 or 5508. http://www.dtic.mil/dtic/tr/fulltext/u2/a801420.pdf  http://airfoiltools.com/airfoil/naca4digit?MNaca4DigitForm%5Bcamber%5D=5&MNaca4DigitForm%5Bposition%5D=50&MNaca4DigitForm%5Bthick%5D=8&MNaca4DigitForm%5BnumPoints%5D=81&MNaca4DigitForm%5BcosSpace%5D=0&MNaca4DigitForm%5BcosSpace%5D=1&MNaca4DigitForm%5BcloseTe%5D=0&yt0=Plot


text:
Why is V1 based only on using brakes?
V1 is the speed that an aircraft may abort a takeoff so long as the airspeed is below it, or is committed to flight if it is over that speed.
V1 calculations are based only on the brakes (not sure about spoilers) being available to stop the airplane (regardless of the number of engines), why?

Doesn’t that depend on certification basis and runway state? In EASA world, V1 for virtually all commercial jets flying these days, on a wet runway, is calculated taking credit for reverse thrust on operating engines. Spoilers are considered on dry and wet runways. All of these devices can be considered inoperable, but then V1 will change (except of course if V1 wasn’t determined by the stopping case to begin with).

Nope, thrust reverse is not allowed to be considered for V1... interesting, I am not sure about spoilers, (I hadn't thought of them) but they probably are allowed.

I’ve always thought it was a conservative distance in that if you forgot to deploy spoilers or thrust reversers you would still be able to stop within the calculated distance.  Airplanes with auto spoilers include data for takeoff distance with and without spoilers.

Before V1, abort. After V1, continue.  Simple, isn't it?
Brakes only during an abort, no thrust reversers.  Why abort? Loss of an engine means the thrust reverser is lost on one side and using the thrust reversor on the other side would cause asymmetrical thrust and might cause loss of directional control, especially on a contaminated runway.

"Before V1, abort. After V1, continue."  Not always (although, fortunately, situations requiring that a takeoff be rejected after V1 are extremely rare).


text:
Are there any situations in which drones may legaly fly higher than 400ft?

FAA FAR 107 states, "The maximum allowable altitude is 400 feet above
  the ground, and higher if your drone remains within 400 feet of a
  structure."

Are there any exceptions to this rule that would allow a drone to be flown at a much higher altitude than 400ft AGL or within 400ft proximity of a tall structure?

source

It tells you right on the page you linked: "unless you’re flying for commercial purposes and have obtained a written permission from the regulating bodies."

It goes to show a little research pays off. See the answer below...

YES, models my be flown to any height by "Academy of Model Aaeronautics" members under guidelines of a "safety program".
In January of 2016 the AMA requested that the FAA clarify the 400-foot issue in writing. The FAA's response recognized AMA’s role as a "community-based organization" and acknowledged the AMA safety program, including allowing flight above 400 feet under appropriate circumstance.
In this letter, dated July 7, 2016, the FAA states:

"…model aircraft may be flow consistently with Section 336 and agency
  guidelines at altitudes above 400 feet when following a
  community-based organization’s safety guidelines."
"Community-based organizations, such as the Academy of Model
  Aeronautics, may establish altitude limitations in their safety
  guidelines that exceed the FAA’s 400 AGL altitude recommendation."

This letter confirms that sailplanes, large model aircraft, turbines, and other model aircraft can legally operate above 400 feet if the AMA member is operating within the AMA safety program. The FAA acknowledges AMA as a community-based organization.

Page 1

Page 2

To add to this answer, the AMA’s guidelines only state that the 400 foot limitation only applies within 3 miles of an airport. “When within three (3) miles of an airport, model aircraft should remain below 400 feet above the ground unless there is notification or an agreement with the airport that allows models to safely go higher.” - From the AMA Safety Handbook

All obsolete now--

As they say with any regulation, there’s a waiver for everything.  If you do wish to operate your sUAS at or above 400’ AGL, you are advised to contact the local FSDO in your area.  Protocol varies from area to area, but you will be required to file an request with local ATC with the dates, times and areas you intend to operate in so NOTAMS can be issued.


text:
What is this pusher airplane?
I am attempting to identify this aircraft.
The pictures appear to have been taken around 1960.

It is the only Bede XBD-2, from the early '60s.
According to Wikipedia it was an experimental aircraft with STOL capabilities.

The tail number must have changed since then, now N327BD is a Piper J3C-65

@ratchetfreak I guess it would be expected if that number is not in use for a certain number of years, no?

The aircraft is on static display at Manitowoc airport and still carries the same tail number, as can been seen here

Who would have guessed the N...BD stands for Bede. I wonder if he did that with his other planes?

@jwzumwalt Seems that it is the case for several prototypes, see photos here of several models


text:
Why does the B-52 outboard engine nacelle have a sharp change in shape?
The outboard engine nacelle of a B-52 appears to have a change of shape too extreme for just blending to the center-line of the engine pod. Why is the B-52 outboard engine nacelle shaped the way it is?

I am not a designer but, two closely spaced intakes competing for air need this type of nacelle to live harmoniously?

The B-52 was built in different versions (A-H) and the engines and their installation differ between these version1.

The image in the question most likely shows the engine configuration of a B-52H.
The following drawings suggest that this also changed the inlet design2.

There seem to be different engine-designations depending on the application. For the B-52H, JT3Ds (TF33-P-3) were used3.
Pictures of the TF33-P-3 inlet show they are symmetrical. (While the question is focused at the outboard engine-inlet, this question can only be answered taking the other engine and the placement relative to the wing into account as well).

At the symmetry-plane the TF33-P-3 inlet-section extends further than the rest of the inlet-lip.
I would guess that there are multiple reasons and considerations which led to this design. Three of which will most likely have been:

Basic aerodynamic considerations like stagnation-point and pressure distribution will prefer a sharp/point nose over a flat area. Especially since the air needs to be diverted from the engine-pylon.
The cruise speed of the B-52H is Mach 0.914 which means the nose could also be some kind of anti-shock-body reducing wave-drag by applying the area-rule.
The nose between the two engines prevents aerodynamic interaction of the two engines, for example during cross-wind operation.

The negative aspects of placing two jet-engines close to each other were also discussed in another SE-question with respect to commercial airliners.

[1]:www.airpowerstrategy.com/
[2]:http://www.afwing.com/
[3]:en.wikipedia.org/wiki/Pratt_%26_Whitney_JT3D
[4]:www.globalsecurity.org/wmd/systems/b-52.htm

A lot of pictures of this intake make it quite hard to determine what is the exact geometry. But the one below is by far the best - thank you @ymb1.

http://www.af.mil/News/Photos/igphoto/2001513002/
Before, it was hard to tell exactly what's happening. In some pictures, it looked like the inboard engine was slightly further forward of the outboard engine.
But that is now shown to be incorrect. The above picture shows where the two intakes touch, the duct protrudes further forward, with a sort of "nose". This can be clearly seen from the parts of the inlet lit by the sun, and the bump in the shadow of the engine on the ground.
I would strongly expect this is to keep the airflow smooth, and reduce aerodynamic interference between the two inlets, and avoid turbulence that would cause inlet pressure distortion. This is, to avoid a difference in total pressure for each engine, between the side that is adjacent to the other engine, and the side that isn't. Uneven total pressure, or inlet distortion, can cause compressor stalls.

Why does the B-52 outboard engine nacelle have a sharp change in shape?

Not just the outboard, but all 8 engines on a B-52 have the same inlet design.
The original engines had a smaller fan diameter and symmetrical round inlets.

I suspect going to a larger fan diameter created flow problems which the engineers solved by modifying the inlets.

Hint: Add something along the line that flow with one engine of a pair out will be less disturbed.

This does not provide an answer to the question. To critique or request clarification from an author, leave a comment below their post. - From Review

@mins I think the answer's fine now that it's been edited (though something more authoritative than "I suspect..." would be even better.)

The original engines didn't merely have a "smaller fan diameter" - they were turbojets, and, as such, had no fan at all.

That looks like an optical illusion due to the angle to me. A photo on this site shows a more frontal view of the outboard engines, with just the equal centre division between the two engines appearing. Close-up underneath.

This inlet design was already used on the Convair B-36 D, which had twin turbojets added on the outer wing to give it a higher top speed when penetrating enemy airspace. During cruise, only the six piston engines would run and the jet engine inlets were plugged to reduce their aerodynamic drag. This plugged state is shown in the picture below (picture source):

Since the B-36 was a purely subsonic design, any intention of area ruling in this design can be confidently ruled out (pun intended).
The same engine was used on the Boeing B-47, but here the twin intakes of the inner nacelle show no such protruding center (below, picture source):

Seems the Convair engineers knew something that the Boeing engineers didn't. But in Seattle they caught up eventually, and the larger airflow of the P&W TF-33 turbofans made them use the same trick from the Boeing B-52 H on. All versions before used the J57 turbojet and the simple B-47 style nacelle shape (B-52 F below, picture source):

The reason is the distorted airflow from spillage if one engine of a pair has to be shut down. By pulling the divider forward in the center of the nacelle, any spillage will occur away from the running engine, ensuring symmetric intake flow there.

Are you sure about the B-47?  The inlet of the outboard engine in the inboard nacelle looks suspiciously like it does, in fact, have that sharp V-shaped notch in the intake...

@Sean: Yes. The question is about the nose between the engines, not the notch. A notch was already pioneered by the Arado 234 V8.

That is a perspective issue. The inlets on all B-52 models are smooth. The picture in the question makes it look like there is a sharp geometry change, but there isn't.


text:
What is the wear limit for tires?
I have 2500hrs and have flown over 30 types of aircraft. I am also a A&P. I have never seen a POH, manufacture or FAA suggestion for wear limits of a tire. The flight school I got my pilots license (1980), would wear the tires of Cessna training aircraft bald on the outside of the tire and flip them on the rims, then run then bald on the opposite side. On airliners and transport aircraft I have seen chunks of rubber missing, cord and belts exposed, etc more so on freight than passenger aircraft. Surely there must be some standard or legal requirement.
( I have seen FAA, mfg, and other info prohibiting flight if tire pressure exceeds +-5%)

I would expect the tire manufacturer would provide a limit.

Yes, there is, but it is not by regulation. It is in Advisory Circular 20-97B.
For example, section 7 would apply to most of your pictures:

(7) In-Service Inspections (Operator).
(a) Tread Damage Removal Recommendations. Tires that exhibit any of the following characteristics symptomatic of damage should be removed from service:

(1) Cracking and/or cutting in a tread groove that is undercutting material adjacent to the crack or cutting that has lifted or can be lifted.

NOTE: Do not probe cuts or embedded FOD when the tire is inflated.

(2) Undercutting in a tread rib.
    (3) Transverse cracking in the tread that is visible at the surface and shows progression along an angular path down into the rubber.

(b) General Tire Removal Recommendations. Tires with any of the following conditions should be removed from service:

(1) Sidewall cuts that penetrate the outermost structural ply.
    (2) Internal separations characterized by bulges or distortions in the tread casing or sidewalls.
    (3) Heat or slippage evidenced by reverted or charred rubber above the wheel flange.
    (4) Tires worn or flat-spotted to the outermost casing ply (bias) or outermost belt (radial).

Well, that is still a "should", not a "must". Even the manufacturers use "should" a lot; the hard limit is when the first ply is worn through, see page 26 of linked PDF.

@PeterKämpf True, not saying that these are "must" items, the OP did ask if "I have never seen a POH, manufacture or FAA suggestion", FAA Aircraft Circulars are pretty much "FAA suggestions". Your linked PDF is for Goodyear tires, it is possible that some other manufacturer includes wear indicators or some other method. Also, see 7.b.4 in the answer "Tires worn or flat-spotted to the outermost casing ply..", which matches your PDF.

The Russians have simple stud holes in the tire surface. If the holes are no longer visible (because all the rubber around them has been worn off) the tires must be replaced. TomMcW is correct, in the West the tire manufacturer and the operator decide what the limits are. You will find similar documentation from Michelin and others. But yes, the question is about suggestions also.


text:
What is the working principle of a Gurney Flap?
Recently I came across a airfoil improvement called a 'Gurney flap', see image.

I don't really understand the working principle of a Gurney.
Wikipedia states:

The Gurney flap increases lift by altering the Kutta condition at the trailing edge. The wake behind the flap is a pair of counter-rotating vortices that are alternately shed in a von Kármán vortex street.In addition to these spanwise vortices shed behind the flap, chordwise vortices shed from in front of the flap become important at high angles of attack.

I don't understand how the flap could generate vortices in front of the flap.
Wikipedia states the following effects:

The Gurney flap increases the maximum lift coefficient ($C_{L_{max}}$), decreases the angle of attack for zero lift ($\alpha_0$), and increases the nosedown pitching moment ($C_M$), which is consistent with an increase in camber of the airfoil. It also typically increases the drag coefficient ($C_d$), especially at low angles of attack, although for thick airfoils, a reduction in drag has been reported.

Is it simply a ridge that stops the flow, and thereby increases the pressure at the top surface?
How come the drag is decreased for thick airfoils if it generates extra vortices?

Perhaps a useful diagram of the Kutta condition: http://people.rit.edu/pnveme/MECE356/lift/lift_images/kutta.jpg

@MikeFoxtrot: The flow behind the wing is wrong on that picture and without description I don't think I would understand anything from it I didn't know already.

I'm not sure Dan Gurney understood why or how it worked, he just knew that he got more downforce on his car when he added that little lip to the rear wing.  ;)  (Yes, the Indy 500 is this weekend, and it's just down the road.)

This does answer a part of my question, but it doesn't talk about the drag advantages of Gurney flaps.

@ROIMaison: NASA tested and simulated the behaviour of Gurney flaps in the past. I find this report in particular interesting. P.s.: you are quite patient, you have waited for 7 years to get an answer :-)

The basic answer is that it creates a low-pressure zone that helps pull air over the upper surface. Thick airfoils often have difficulty maintaining enough energy in the aft part of the upper surface. The low energy boundary layer becomes thick and very prone to forming a turbulent bubble. (aka a stall or detached flow)
At the same time, it slows the flow on the lower surface getting resulting in increased static pressure.
The reduction in drag is due to better control over the size and placement of the detached upper flow.
Also be careful to compare apples to apples. The fixed AOA on a car is not equivalent to the variable AOA of an aircraft. There are considerations like racing rules that limit dimensions, weight, and cost the Gurney flap was a much cheaper modification to that situation than a whole new wing, stall characteristics, stiffness and weight requirements of the overall wing shape, and available materials.
As an aside the x-15 tail has pure wedges for airfoils, just flat-sided triangles, pointy front and flat back. But that's a hypersonic deal, more like deflecting particles than any pressure-induced flow patterns.

The Gurney flap diagram illustrates a recovery of air flow energy that is a long known mechanism for improved L/D ratios of wings with drooped leading edges:  a vortex returning energy back into the wing.
Notice the diagram in the question has three of them.  The undercambered wing would have one large one underneath.
This device works by increasing pressure "underneath" (opposite lifting force) and smoothing the airflow on "top", much the same way as a normal flap does: by increasing trailing edge downwash.
With the Gurney flap, airflow off the top and bottom power the two trailing vorticies, which return some of their energy to the flat surface of the flap.
But these benefits may exist only at a certain angle of attack.  More research would be in order.
However, as seen with Krueger flaps, air flow, as well as physical form, ultimately determines Lift/Drag properties.


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Why do T- tail airplanes have a shorter vertical stabilizer?
Airplanes with a T-tail design have a shorter vertical stabilizer than the conventional design , what is the reason?

related: https://aviation.stackexchange.com/q/1400/1467

Besides what's been said in answers, seems to me that if you had to deal with loads generated by horizontal stab  on top of vertical stab, you'd have an incentive to use a shorter, broader vertical stab.

Two reasons:

T-tail design is often imposed on designs with twin engines mounted at the aft fuselage. This means they have a small moment arm in the yaw direction, the vertical tail is dimensioned to compensate for engine failure. The further away form centreline the engine is mounted, the higher the yawing moment that the remaining engine exerts upon engine failure.
The horizontal stabiliser at the top of the vertical stab functions as an aerodynamic barrier, like a winglet does. This prevents a vortex from streaming off of the top of the vertical tail when yawing, and thereby increases the effectivity of the vertical tail surface. This effect was researched in NACA TN 2907, at five different locations and three different spans:

The graph below shows calculated effective increase of aerodynamic aspect ratio, as function of where on the fin the horizontal tail is located. If completely at the top or at the bottom, the aerodynamic aspect ratio is about 50% higher than when at the centre, resulting in increase of the lift curve slope of the fin of roughly 15%.
With a T-tail, the horizontal stabiliser at the top functions as an aerodynamic barrier, while the fuselage at the bottom end provides some barrier function as well. Depending on the local fuselage circumference and shape of course.

The last plot does not seem to show what you say. It appears that placing the wing at station A (bottom) or E (top) give the same effective aspect ratio, both 50% better than placing at the midway up the vertical stabilizer. This makes sense to me -- shouldn't the endplate effect work on either end of a lifting surface?

@supergra Good catch! Have amended.

Does the first bullet point reverse cause and effect? I thought it's not so much that a T-tail is imposed when you have fuselage-mounted engines, bit rather that it necessitates a shorter tail, which in turn requires fuselage-mounted engines (so that the moment is lower). Or alternatively, That it's enabled (not imposed) by that engine configuration.

Another reason according to Wikipedia is to reduce inertial coupling. I haven't understood so far how that works – can you shed some light on the issue?

Raymer gives the following answer:

The 'T-Tail' is also widely used. A T-tail is inherently heavier than
a conventional tail because the vertical tail must be strengthened to
support the horizontal tail, but the T -tail provides compensating
advantages in many cases.
Due to end-plate effect, the T-tail allows a smaller vertical tail.
The T-tail lifts the horizontal tail clear of the wing wake and
propwash, which makes it more efficient and hence allows reducing its
size. This also reduces buffet on the horizontal tail, which reduces
fatigue for both the structure and the pilot." (pg. 69)

- D. P. Raymer, 'Aircraft Design: A Conceptual Approach', AIAA, Washington DC, USA, 1992.


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Why wasn't the 737 MAX's landing gear lengthened enough to mount the engines under the wings, rather than in front of them?
The reason Boeing originally added the MCAS to the 737 MAX was to compensate for a slight pitch-up moment generated by the engine nacelles at very high angles of attack.  This pitch-up moment occurred because the LEAP engine used on the MAX series is larger and longer than the CFM56 used on the 737 Classic/NG, so Boeing mounted the new engines further forward and higher up to fit them on the plane; the larger, forwarder nacelles added more planform area to the front part of the aircraft, causing the MAX's center of lift at high attack angles to be slightly forwarder than on the Classic or NG, causing the MAX to pitch up slightly more at high attack angles.
However, the reason the Classic and NG have their engines mounted in front of the wing is that the 737 originally used the much slimmer JT8D engine, allowing it to use a very short landing gear to make it sit lower on the ground and ease the process of loading baggage and other cargo by hand.  When the Classics came along, Boeing decided to keep the 737-100/-200-length landing gear, which forced them to use a smaller-fan variant of the CFM56, move all the engine accessories from the bottom to the sides, and mount it in front of the wing (instead of directly below it) in order to give the new, larger engine sufficient ground clearance; with the NGs, which use a CFM56 variant that is essentially the same size as that used on the Classics (with a diameter all of one inch greater), they simply kept the same nacelle, pylon, and landing gear designs as those used on the Classics.
The MAX, in contrast, has a longer landing gear than all other 737s (handloading of baggage not really being that much of a thing anymore at most 737-served airports); given that they were giving the MAX a longer landing gear anyway, why not make it long enough to let the engines fit under the wings and finally eliminate the engine-mounting kludge inherited all the way from the 737-300?

I would have thought you'd understand the design modification merry-go-round by now Sean.

The MAX 10 has a new landing gear because of the higher tailstrike risk during rotation. It shrinks during retraction to make it fit into the small landing gear bay, which makes it quite complex (more complexity = more maintenance cost).

Considering how hard they cheaped out on the MACS, cost is the most probable reason. It doesn't have to be prohibitively expensive to redesign the landing gear or the landing gear bay. Just a little more expensive is enough. They thought they could get away with the new engine mount, they didn't, hence MACS. They thought they could get away with MACS, they didn't, hence the 300+ lives. Bottom line is a bitch, so is duct tape engineering.

Interesting how all the answers talk about the mechanical problem of fitting the wheel. I was thinking from a human performance pov, it would be a nightmare to maintain a common type rating for pilots, changing flare height everyday depending on type flown.

@quietflyer - Well yes, but the probability of surviving is inversely proportional to the flight path angle

The main gear sits between the wings and retracts inwards, such that the wheels will rest in the lower fuselage. Lengthening the main gear would either require to shift the gear attachment points outward or to retract one leg a bit forward and the other a bit backwards, such that both wheels will sit behind each other instead of side-by-side, as they do now - see below (source).

Shifting the gear attachment points out is clearly feasible but results in a major wing redesign. That would cost money so Boeing management decided that a software fix would be the better solution.
Placing the wheels in tandem would equally be a major and costly modification. A lot of systems would need to be relocated and all modifications would require testing. Boeing management cleary did not want to pursue this option.
Another quick fix is shortening the gear upon retraction. Decompressing the oleo strut in order to shorten the gear upon retraction is already standard in many fighter jets. Another solution is to add another hinge such that the gear leg will fold inside the gear bay. Such a solution was used on the Grumman F-8F Bearcat and is pictured below (source).

From what I understand, this particular solution was applied to the longest 737 version to allow enough tail clearance for rotation, but added weight and complexity. Of course it could be applied to the rest of the 737 MAX fleet as well, but again Boeing management tried to save money and weight by implementing a software-only solution.
But the engine location would be fine if a larger horizontal tail would be added, increasing stability at high angle of attack and making a quick software fix unnecessary. Again, this would add cost and weight and make the aircraft less competitive.

For further reading on Boeing's ill-fated  descisions regarding development of 737 MAX, please google "dead horse theory".

Not so long ago, Boeing released very short video about MAX 10 main landing gear. As I understood it, the main focus was tail strike avoidance and it looked like the ground clearance (during taxi, etc) was the same as with MAX 9/8. So perhaps the CFM LEAP engines still had to be relocated where they are now.

Because lengthening the main gear struts constitutes an almost complete structural re-design of the main wing structure, with dramatic consequences on cost, lead time, re-certification, weight increase etc.

The main landing gear of most swept back winged aeroplanes retract behind the rear spars, in the kinked section close to the fuselage. Picture above is also displayed in this answer, and shows the main gear hinge point for the DC8 in highlight. One of the issues with lengthening the gear can be seen immediately: the hinge point must be located further outboard, interfering with the flaps/aileron mechanisms. The wheels retract as close as they can to the main centre line, as seen below in a 737 bottom view.
Image source

But the main issue is with absorption of the landing impact forces. The main gear is dimensioned to withstand a certain vertical speed at max. landing weight, resulting in an impact force F. If the distance from centre is increased, the impact torque T on the centre wing box increases proportionally. Strengthening the centre wing box has implications on the weight and/or the wing root thickness, which has implications on the landing gear impact force, which is part of a design merry-go-round that needs to be checked with construction departments, aerodynamicists, etc etc.

Torsion-wise, the same thing happens. The main gear is mounted behind the torsion box, and the landing impact force must be compensated for by torsion in the section between main gear and wing root. Mount the main gear further outboard, and a larger moment arm acts on a smaller local cross-section: more strength and weight and structures and design procedures required.
Image source, close-up
Above cut-away drawing shows the wing-fuselage intersection of a B737. The wing centre section is the critical part in the support structure, therefore it is not interrupted anywhere and runs through a cut-out in the fuselage, which then needs to be re-inforced at the front and rear spar pick-ups. The most complicated part of the aeroplane - any changes in load will have far reaching effects.
In the initial design phase, any length landing gear will be incorporated in the design merry-go-round - it will take a long time before the type certificate is in place, regardless of the main gear length. But once the design is fixed and in place, a redesign of the fuselage/wing section is a major effort. So if an aeroplane factory is under competitive pressure to release an updated design, they will want to leave the wing structure unchanged as much as possible.

If the wing is redesigned anyway for a larger span, lengthening the main gear struts will be part of an integral major design effort.
If wing structural design is kept relatively unchanged when releasing a version upgrade, then almost by definition the main gear strut length must be kept constant.

As you describe it, the B757 or the A320 are technical miracles, having solved a further outboard gear attachment point problem that has stymied Boeing in the 737. No, the answer is more simple: They wanted to avoid the extra cost and weight of a longer gear. They had plenty of chances of getting it right and other designs show that it is a straightforward thing to do.

@PeterKämpf No. The B757 and A320 were not redesigned for a longer landing gear. Redesigning the wing/fuselage intersection is a major design effort in cost and delivery time, as any aeronautical engineer worth their salt knows. Accommodating a longer gear in the initial design process is just one variable in the aerodynamic/contstructive design matrix.

The rear wing spar is unswept inside of the engine attachment points. Reinforcing it further outboard is not rocket science. Of course it costs more than having 10 $ per hour programmers hack up a software fix and certifying that with in-house people. But it is far from impossible.

@Koyovis your appreciation of the complexities of the design and difficulties changing it have merit, but professional engineers earn their "salt" by successfully dealing with these challenges.  1.  They should have seen larger fans coming.  They had years to lengthen the gear FIRST.  2.  If they can't make the wing 2 feet longer, the other option (obviously) is the make the fuse longer and put the engines there.  A dedicated engine mount section would do it, balanced with another fore of the wing.

YES anything is possible. As you can see, I'm explaining why it is a major redesign issue. I'm not claiming that I know better than the Boeing company, that would be arrogant.

Well, the ones sitting in the parking lots could also have the NG engines put back on and sold at a discount.  The delayed Y1 "cleansheet" will hopefully take care of all of this, lessons learned.  They certainly have their work cut out for them.

And a Y1 proposal:  have a look at the B-47 MTOW and it's overall design.  Yes, pretty close to the 737!  Idea began as a hilarious muse considering the results of retired B-47 engineers showing up to help, but ... ?

Given the technical competence of Boeing management, the proposed claim of you to know better is not arrogant but borders on the trivial.

@PeterKämpf Getting pretty tired of some people who know everything. How did our ex-prime minister put it again?

@PeterKämpf Is Boeing management really technically competent? As I understand it, a lot of engineering-first culture was lost when the bean counters from the McDonnell-Douglas acquisition ended up in positions of influence.

@T.J.L. That is exactly right. Boeing has many capable engineers (well, fewer than they used to have, but anyway), but the usual ladder into top management was cut with the McDonnell-Douglas acquisition. Now everything is decided on a cost basis with no regard for long-time consequences.

Unfortunately the short and simple answer is money. They weren't prepared to spend the money on a costly redesign of the already proven 737NG fuselage.
The best in depth explanation I've seen of the whole issue with the MAX is: "How the Boeing 737 MAX disaster looks to a software developer".

Thanks for that article. I did not know that the computers only monitored their own sensors and not the ones on the other side.  You are right, that was an gross oversight.

+1 for the article.

The technical answers above are fairly good, but probably the biggest answer is that Boeing corporate out in Chicago was simply too miserly to invest in a clean sheet airplane (which it could have EASILY done) and demanded that Boeing Commercial Airplane develop this new medium range airliner on a limited budget within a limited time span.  The 737 MAX simply shows how far Boeing was willing to try to stretch an old design to make it work with modern engines and modern technology. In this case it was simply pushed beyond its limits - and definitely beyond anything that it’s designers in the 1960s could have envisioned the 737 being used for.

There is a limit to what can change on an air-frame before being considered a new aircraft.  I mean new as in a whole new designation.  That means a whole new approval cycle that may or may not be approved by the FAA or international organizations.
Moving a pivot point is not a simple task, there is a lot of structure to move. If you move structure there is a lot of systems lines and wires that need to move.  Because you have moved all this has moved there interference checks need to be made to make sure that nothing is touching that should not touch.  Weight and balance then needs to be checked to make sure that it is still in limits of the aircraft.  We started this to make room for a longer gear, but now the gear doesn't meet the side load requirements because its longer and in a different spot.
  I am willing to bet that some where in this process, the guidelines of what I talked about above have been met,  suddenly your costs have sky rocketed.
Do you have any experience with CADD software?  If you do then you know the problem if a part in your assembly suddenly changes size.
Did Boeing go the cheap way to fix the MAX?  Probably
Did they find a good way to fix the problem?  I think so
Did they implement this fix correctly?  Don't think so
Should this have become required known information? Yes

Boeing's main problem was time to market - they were in a tremendous hurry to compete with the latest Airbus. Re-engineering the undercarriage would have been more complex and taken a lot longer than adjusting the design of the new engine pylons.
Perhaps ironically, the latest Max 10 has had the time to develop an extending undercarriage, giving it 9 inches (23 cm) extra ground clearance. Although the Max 10 needs that to keep its longer tail off the ground, it would have helped a lot with the engine mounting problem too, had Boeing introduced it on the first Max. There might then have been no requirement for MCAS.

The fuselage is not wide enough and therefore there is not enough space in the wheel well to store two main gears, had they been extended further.
MAX 10 employs a new landing gear design that shrinks the legs slightly, but even that is just on the limit of what is possible, so if Boeing increased engine size for all models and used that mechanism for MAX 7, 8 and 9, they would have been unable to develop the MAX 10.

this does not make sense. The MAX series uses basically the same engines. How would adopting MAX 10 ldg gear on 7, 8, 9 make it impossible to develope MAX 10?

Had they used the MAX 10 landing gear for MAX 7, 8 and 9, there would be no further room for extending the landing gear in order to accomodate MAX 10, which requires even longer landing gear than the other three models to ensure sufficient tail clearance during rotation.

If MAX 10 ldg gear was used in 7/8/9, why on earth would the 10 then need even longer struts? They were already extended once to give tail clearance, where would this requirement for a second extension come from?

The first sentence does not make sense. If the wheel base would be extended, the mounting points would move outwards, but the wheels wells would remain in the same place.


text:
Which aircraft has the highest minimum operating temperature?
Aircraft often have a minimum operating temperature due to fuel, oil, engine, and other limitations.  Which aircraft has the highest minimum operating temperature, and why?
I'm looking for something along the lines of:

A Cessna 152 has a minimum operating temp of -5°F while all other aircraft can operate at -6°F or colder, therefore the 152 has the warmest "minimum operating temperature" of all aircraft.
Note: numbers made up for an example

Clarification based upon comments below

Don't you mean lowest minimum operating temperature? Most transports are cert to -40C, or sometimes -45C.

@JohnK, no I do not.  I would guess some aircraft have water in them and cannot be operated below 0C.  Maybe there are some with higher lower limits?

Potable water for the Lav system will be loaded with everything running and the interior up to temperature, plus there will be heaters at key points to prevent freezing. Airliners dispatch at temps down to -40 every day in the arctic. Certified minimum operating temps are typically -40C, which means below that, you have to wait for the day to warm up to -40. All the components on the airplane will have been qualified in development to function at -40. In the air, hardware not within the pressure hull may be expected to work down to -60.

@JohnK Yes, but none of that has anything to do with the water ballast in a sailplane.  The question is broader than airliners.

Sailplanes don't have fuel, oil ,engine and other limitations lol.

If you're interested in sailplanes, you may want to [edit] your question to specify that. If there's something else you're after, an edit would certainly help clarify that, too. As it stands, though, I think you're looking for something like "A Cessna 152 has a minimum operating temp of -5°F, therefore it's got the highest minimum operating temp." (number completely made up) Is that what you're looking for?

Wildfires need a relatively high temperature to ignite, so firefighting aircrafts have a quite high minimum operating temperature

@FreeMan Yes.  It is not sailplane specific.

I think that would refer more to the maximum operating temp, @sophit. Or, if you're thinking that a fire might only start in the summer when the temp is 95°F/35°C, then that's not the minimum operating temp, that's just the current operating temp.

@FreeMan obviously I was just joking. But yes, since under 35°C firefighting aircrafts are not operated then that would be the minimum operating temperature


text:
Why does my Surface Analysis chart indicates radar returns / rain when none exist?
I'm using 1800wxbrief.com to get all the information necessary to brief myself on an upcoming flight.
I have noticed each time I've done this, that the Surface Analysis chart shows light / spotted rain where I can be nearly positive there is none. When cross checking this chart with sites like weather.com, it is confirmed through their weather products that there is no rain in the area.
Do these Surface Analysis charts sometimes show spotty false positives?

I believe this is an artifact around the actual radar receiver, usually you see this right around the radar itself.

@RonBeyer -- yes, NWS radar products (weather.gov) sometimes show clutter artifacts when the underlying NEXRAD system kicks into high gain (i.e. clear air) mode.

Some of those will be flocks of birds.

Fog and other low-altitude moisture often shows as very light precipitation radar returns. I look at the multi-image radar loops for an area, and if I see that I call a "sparkling" radar return in the looping image, it's a pretty good bet for ground fog. The California central valley area (around Modesto and Stockton) is known for its winter tule fog: watch the radar returns from that area in the wintertime.
You can correlate the radar observations with ground visibility observations (airport METARs, the regular and special airport weather observations, include visibility and ceiling—cloud height and coverage—information), along with satellite visible and infrared images to view cloud presence and cloud top heights. When the visible satellite shows clouds and the infrared satellite shows very low cloud tops, that's a good bet for fog or a low-level stratus layer: both have low cloud tops, typically 400'-1,000' above ground level.

Thank you for the great answer @ammPilot, and welcome to the site.


text:
How is it possible to perform a rapid climb / tight loop without stalling?
I'm still trying to get my mind fully around "Angle of Attack". This makes sense to me in the most basic situations, but when I start throwing more dramatic situations at it, my understanding breaks down... which means I don't really understand it.
Take this video for example (at  0:17 and 1:32):

How is this possible? Is it due to the excess power of the aircraft allowing it to change its "flight path" constantly through the maneuver and thus keeping the flight path (relative wind) not far away from the Critical Angle of Attack?

Taken to an extreme, a helicopter can hover with 0 relative wind because it has enough power.  An overpowered airplane is similar and can "hang" on the prop in much the same way.  It looks like most of the aerobatic maneuvers in the video though are using intertia and trading airspeed for altitude and are not sustainable because they will eventually run out of speed and stall. (I'll leave a more technical answer to someone who has the time to provide one.)

@Lnafziger: And to take things to the other extreme, it's perfectly possible to perform steep climbs (for a short while), loops and other aerobatic maneuvers in a sailplane.

Can you link to the video not on facebook?

Does http://aviation.stackexchange.com/q/2903/524 clear up your confusion?

change its "flight path" constantly through the maneuver and thus keeping the flight path—the term "flight path" is used twice in a rather inconsistent way. Can you try to reword it to clear it up?

Short answer:
The maximum angle of attack is never reached in this video, thus the aircraft is not stalled.
Longer answer:
Stall is a problem mainly occurring during low speed flight. Our angle of attack is always depending on our flight path.
Let's first assume, that are at cruising altitude in a level flight condition. In this case the pitch angle of the airplane equals our angle of attack. As generated lift is dependent on your airspeed and the lift coefficient (which is again dependent on your angle of attack), reducing airspeed in level flight while maintaining your altitude will force you to in increase the angle of attack. At some point a further increase will result in a too high angle of attack and thus stall the aircraft.
Now, let's look at your problem:

(source: aeroskytech.com)
The main difference is your speed, or so to say the excess power used for climbing. During climb, the trajectory of the plane is not equal the horizontal axis. Therefore, also your angle of attack is not equal to your pitch angle (the angle between the longitudinal axis of the plane an the horizontal axis), but to the angle between the trajectory and the longitudinal axis. During the loops showed in the video the airplane is not only changing its pitch angle but also its flight path, therefore it is not stalled.
An interesting example for maintaining your flight path while increasing are military aircraft in combat maneuvers, their pilots rapidly change their pitch angles while still flying in the same direction. This works as a decent speedbrake, allowing them to intercept other aircraft.

What causes the mist on the picture above? AFAIK, it is water vapor condensing due to sharp fall of the pressure and therefore - temperature - over top of wings. Am I right?

Drastic drop in pressure over the top of the aircraft, which in turn results in a drop in temperature and thus the air flowing through that section is at a temperature below the dewpoint of the air.  Moisture vapor forms as a result.

The example of the GB-1 going vertical at 0:17 in the video will work just fine, so long as the pilot does not exceed the critical angle of attack on the wing. It's being flown by a skilled aerobatic pilot who is familiar with the GB-1 flight envelope and is a pretty benign maneuver. For an aircraft like that, I'd guess the maneuver begins around 160-170kts at a load factor of 4-6Gs; a decent pull on the stick but not enough to reach the critical angle of attack for those airspeeds. In addition, the stall characteristics of those aerobatic aircraft are pretty benign; the onset just feels like buffeting and shuddering in the airframe and can be relieved simply by easing off stick pressure.
The example at 1:32 is a post stall maneuver where the plane is stalled and is simply hanging on the prop.  You can do this if you have enough power, and the maneuver begins at or near Vs.

Excess power, yes! But also light weight and a robust, low aspect ratio wing.
Excess power helps maintain adequate airspeed at all attitudes, including straight up.
Prop wash also maintains airflow over the wings and particularly the empenneage control surfaces, but torque effects must be managed at low airspeeds.
Light weight is prized by aerobatic pilots because the aircraft can change directions easier and "follow" the pitch change, rather than continuing in the same direction while AOA increases.  This is a major problem with very heavy aircraft that have high wing loading such as airliners.
The other side of stalling due to excessive AOA is exceeding the G force (load) limits
by pulling too hard on the elevator at excessive speed.  The wings break before you reach stall AOA.  Here the strength of the low aspect wing saves you, and the G forces serve as warning.
Finally, the low aspect wing stalls at a much higher AOA and lower airspeed than a high aspect wing.  Have a look at what most modern fighter planes have.
Add in a pilot properly trained and experienced in that type of aircraft, and you have one amazing video.

303 hp, 2200 lbs, g load limits +/- 10 (ten)